US20100150705A1 - Gas turbine engine - Google Patents
Gas turbine engine Download PDFInfo
- Publication number
- US20100150705A1 US20100150705A1 US12/591,260 US59126009A US2010150705A1 US 20100150705 A1 US20100150705 A1 US 20100150705A1 US 59126009 A US59126009 A US 59126009A US 2010150705 A1 US2010150705 A1 US 2010150705A1
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- US
- United States
- Prior art keywords
- vanes
- gas turbine
- pressure
- turbine engine
- rise
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R7/00—Intermittent or explosive combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D1/00—Non-positive-displacement machines or engines, e.g. steam turbines
- F01D1/02—Non-positive-displacement machines or engines, e.g. steam turbines with stationary working-fluid guiding means and bladed or like rotor, e.g. multi-bladed impulse steam turbines
- F01D1/023—Non-positive-displacement machines or engines, e.g. steam turbines with stationary working-fluid guiding means and bladed or like rotor, e.g. multi-bladed impulse steam turbines the working-fluid being divided into several separate flows ; several separate fluid flows being united in a single flow; the machine or engine having provision for two or more different possible fluid flow paths
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C5/00—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
Definitions
- the present invention relates to a gas turbine engine and in particular to a gas turbine engine having a pressure-rise combustor.
- a pressure-rise combustor is also known as a pressure-gain combustor.
- a pressure-rise combustor produces an unsteady flow of exhaust gases.
- a pressure-rise combustor produces a flow of exhaust gases which has a large amplitude of unsteadiness. In order to extract mechanical power from the exhaust gases leaving the pressure-rise combustor it is necessary to rectify the flow.
- a problem with this arrangement is that the mixing process occurring in the ejector is an irreversible process, which has a negative impact on the overall efficiency of the pressure-rise combustor. In our tests it has been estimated that approximately 40% of the mechanical energy content of the unsteady exhaust gas flow from the pressure-rise combustor is wasted. In addition the use of an ejector in flow series between the pressure-rise combustor and the stage of turbine nozzle guide vanes considerably increases the axial length of the arrangement. A gas turbine engine with such an arrangement would require longer and heavier shafts than currently used in gas turbine engines.
- the present invention seeks to provide a novel gas turbine engine having a pressure-rise combustor, which reduces, preferably overcomes, the above mentioned problem.
- the present invention provides a gas turbine engine having at least one pressure-rise combustor, the at least one pressure-rise combustor being positioned upstream of a stage of turbine nozzle guide vanes, the vanes of the stage of turbine nozzle guide vanes forming an ejector, wherein each of the vanes having an upstream portion and a downstream portion, the upstream portions of the vanes having leading edges and the upstream portions being arranged substantially straight and parallel to define constant area mixing passages between adjacent vanes for a flow of gases there-through, the downstream portions of the vanes being arranged at an angle to the upstream portions of the vanes to turn the flow of gases there-through.
- the at least one pressure-rise combustor is a pulsejet combustor, a valve-less pulsejet combustor or a pulse detonation combustor.
- a compressor upstream of the at least one pressure-rise combustor and a turbine downstream of the stage of turbine nozzle guide vanes.
- the pressure-rise combustors are arranged circumferentially around an axis of the gas turbine engine.
- the ratio of the distance between the outlet of the at least one pressure-rise combustor and the leading edges of the upstream portions of the vanes to the diameter of the outlet of the at least one pressure-rise combustor is 1 to 1 or less than 1 to 1.
- the turbine nozzle guide vanes have a contraction ratio
- the contraction ratio is the ratio of flow area into the turbine nozzle guide vanes to the flow area out of the turbine nozzle guide vanes and the contraction ratio is in the range of 8 to 1 to 30 to 1.
- splitter vanes There may be one or more splitter vanes positioned between adjacent vanes of the stage of turbine nozzle guide vanes, the one or more splitter vanes are positioned between the downstream portions of adjacent vanes.
- FIG. 1 shows a gas turbine engine having a pressure-rise combustor and nozzle guide vane arrangement according to the present invention.
- FIG. 2 is an enlarged schematic of a pressure-rise combustor and nozzle guide vane arrangement according to the present invention.
- FIG. 3 is an enlarged schematic of a further pressure-rise combustor and nozzle guide vane arrangement according to the present invention.
- a gas turbine engine 10 as shown in FIG. 1 , comprises in axial flow series an inlet 12 , a low pressure compressor 14 , a high pressure compressor 16 , a combustion section 18 , a high pressure turbine 20 , a low pressure turbine 22 and an exhaust 24 .
- the low pressure turbine 22 is arranged to drive the low pressure compressor 14 via a shaft 26 and the high pressure turbine 20 is arranged to drive the high pressure compressor 16 via a shaft 28 .
- the combustion section 18 comprises at least one pressure-rise combustor 30 and the at least one pressure-rise combustor 30 is positioned upstream of a stage of turbine nozzle guide vanes 32 .
- the gas turbine engine 10 is generally conventional and it operates in a conventional manner.
- the vanes 33 of the stage of turbine nozzle guide vanes 32 form an ejector, shown more clearly in FIG. 2 , wherein each of the vanes 33 of the stage of turbine nozzle guide vanes 32 has an upstream portion 34 and a downstream portion 36 .
- the upstream portions 34 of the vanes 33 have leading edges 38 and the upstream portions 34 of the vanes 33 are arranged to extend substantially in straight lines.
- the upstream portions 34 of the vanes 33 are arranged substantially parallel to each other to define constant area mixing passages 40 for a flow of gases there-through.
- the downstream portions 36 of the vanes 33 are arranged at an angle ⁇ to the upstream portions 34 of the vanes 33 to turn the flow of gases there-through.
- the low pressure compressor 14 and high pressure compressor 16 are upstream of the at least one pressure-rise combustor 30 and the high pressure turbine 20 and low pressure turbine 22 are downstream of the stage of turbine nozzle guide vanes 32 .
- the at least one pressure-rise combustor 30 may be a pulsejet combustor, a valve-less pulsejet combustor or a pulse detonation combustor.
- pressure-rise combustors 30 positioned upstream of the stage of turbine nozzle guide vanes 32 .
- the pressure-rise combustors 30 are arranged circumferentially around an axis X-X of rotation of the gas turbine engine 10 .
- the pressure-rise combustors 30 have inlets 42 at their upstream end to receive air from the high pressure compressor 16 and outlets 44 at their downstream ends to discharge exhaust gases towards the turbine nozzle guide vanes 32 and high pressure turbine 20 .
- the outlets 44 of the pressure-rise combustors 30 have a diameter D and are spaced from the leading edges 38 of the turbine nozzle guide vanes 32 by a distance L.
- the ratio of the distance L between the outlets 44 of the at least one pressure-rise combustor 30 and the leading edges 38 of the upstream portions 34 of the vanes 33 of the turbine nozzle guide vanes 32 to the diameter D of the outlet 44 of the at least one pressure-rise combustor 30 is 1 to 1 or less than 1 to 1.
- the turbine nozzle guide vanes 32 have a contraction ratio in the range of 8 to 1 to 30 to 1.
- the contraction ratio of the turbine nozzle guide vanes 32 is the ratio of flow area A I into the turbine nozzle guide vanes 32 to the flow area A O out of the turbine nozzle guide vanes 32 .
- the distance L from the outlet 44 of the pressure-rise combustor 30 to the leading edge 38 of the vanes 32 is maintained small to retain maximum mechanical energy in the flow of exhaust gases from the pressure-rise combustors 30 .
- the upstream portions 34 of the vanes 33 again define constant area mixing passages 40 and each constant area mixing passage 40 supplies a flow of gases into passages 48 between the downstream portions 36 of the vanes 33 and the splitter vanes 46 .
- the splitter vanes 46 are located equi-distant between the downstream portions 36 of the vanes 33 and the splitter vanes 46 have substantially the same profile as the downstream portions 36 of the vanes 33 .
- the leading edges 50 of the splitter vanes 46 may extend in an upstream direction slightly into the mixing passage 40 between the upstream portions 34 of the vanes 33 . If there is more than one splitter vane 46 between adjacent vanes 33 , then some of the passages 48 are defined between two splitter vanes 46 . In this case the distances between adjacent splitter vanes 46 and the distances between splitter vanes 46 and the downstream portions 36 of the vanes 33 are equal.
- the ratio of the number of splitter vanes to pressure-rise combustors 30 may be increased by providing a plurality of splitter vanes 46 between the downstream portions 36 of adjacent vanes 32 .
- the leading edge 38 of the upstream portions 34 of the vanes 32 may have a part-circular cross-section or a square cross-section.
- the cross-sectional area of the outlet 44 of the pressure-rise combustor 30 is less that the cross-sectional area of the inlet to the mixing ducts 40 and preferably the ratio of the cross-sectional area of the inlet to the mixing passage to the cross-sectional area of the outlet of the pressure-rise combustor is 5.5 to 1.
- Adjacent pressure-rise combustors may be arranged to fire in phase or out of phase.
- the present invention provides an axially compact arrangement for a pressure-rise, or pressure-gain, combustor compared with the previous arrangements with an ejector in flow series between the pressure-rise combustor and a stage of turbine nozzle guide vanes.
- the present invention enables the axial length of a gas turbine engine having a pressure-rise combustor to be reduced, reducing the axial length of shafts and hence the weight and cost of the gas turbine engine.
- the present invention reduces the amount of material exposed to high temperature exhaust gases and thus reducing the cooling requirements.
- the exhaust gases from the pressure-rise combustor are mixed with the ambient gas around the pressure-rise combustor in an accelerating region within the passages between the vanes of the stage of turbine nozzle guide vanes and therefore entropy production, or irreversibilities, are minimised, thereby increasing the efficiency of the pressure-rise combustor.
- the present invention produces a steady flow of gases at the outlet of the turbine nozzle guide vanes from the unsteady flow of exhaust gases from the outlets of the pressure-rise combustors.
- the gas turbine engine may be a turbofan gas turbine engine, a turbopropeller gas turbine engine, a turboshaft gas turbine engine or a turbojet gas turbine engine.
Abstract
Description
- The present invention relates to a gas turbine engine and in particular to a gas turbine engine having a pressure-rise combustor. A pressure-rise combustor is also known as a pressure-gain combustor.
- A pressure-rise combustor produces an unsteady flow of exhaust gases. A pressure-rise combustor produces a flow of exhaust gases which has a large amplitude of unsteadiness. In order to extract mechanical power from the exhaust gases leaving the pressure-rise combustor it is necessary to rectify the flow.
- It is known to rectify the flow of exhaust gases from a pressure-rise combustor by providing an ejector downstream of the pressure-rise combustor and upstream of a stage of turbine nozzle guide vanes. The ejector positioned downstream of the pressure-rise combustor produces a pressure field, which causes ambient air/gas to be drawn into the ejector with the exhaust gases from pressure-rise combustor. The ambient air/gas and exhaust gases from the pressure-rise combustor mix together in the ejector and leave the ejector with an approximately steady flow.
- A problem with this arrangement is that the mixing process occurring in the ejector is an irreversible process, which has a negative impact on the overall efficiency of the pressure-rise combustor. In our tests it has been estimated that approximately 40% of the mechanical energy content of the unsteady exhaust gas flow from the pressure-rise combustor is wasted. In addition the use of an ejector in flow series between the pressure-rise combustor and the stage of turbine nozzle guide vanes considerably increases the axial length of the arrangement. A gas turbine engine with such an arrangement would require longer and heavier shafts than currently used in gas turbine engines.
- Accordingly the present invention seeks to provide a novel gas turbine engine having a pressure-rise combustor, which reduces, preferably overcomes, the above mentioned problem.
- Accordingly the present invention provides a gas turbine engine having at least one pressure-rise combustor, the at least one pressure-rise combustor being positioned upstream of a stage of turbine nozzle guide vanes, the vanes of the stage of turbine nozzle guide vanes forming an ejector, wherein each of the vanes having an upstream portion and a downstream portion, the upstream portions of the vanes having leading edges and the upstream portions being arranged substantially straight and parallel to define constant area mixing passages between adjacent vanes for a flow of gases there-through, the downstream portions of the vanes being arranged at an angle to the upstream portions of the vanes to turn the flow of gases there-through.
- Preferably the at least one pressure-rise combustor is a pulsejet combustor, a valve-less pulsejet combustor or a pulse detonation combustor.
- Preferably there is a compressor upstream of the at least one pressure-rise combustor and a turbine downstream of the stage of turbine nozzle guide vanes.
- Preferably there are a plurality of pressure-rise combustors positioned upstream of the stage of turbine nozzle guide vanes.
- Preferably the pressure-rise combustors are arranged circumferentially around an axis of the gas turbine engine.
- Preferably the ratio of the distance between the outlet of the at least one pressure-rise combustor and the leading edges of the upstream portions of the vanes to the diameter of the outlet of the at least one pressure-rise combustor is 1 to 1 or less than 1 to 1.
- Preferably the turbine nozzle guide vanes have a contraction ratio, the contraction ratio is the ratio of flow area into the turbine nozzle guide vanes to the flow area out of the turbine nozzle guide vanes and the contraction ratio is in the range of 8 to 1 to 30 to 1.
- There may be one or more splitter vanes positioned between adjacent vanes of the stage of turbine nozzle guide vanes, the one or more splitter vanes are positioned between the downstream portions of adjacent vanes.
- The present invention will be more fully described by way of example with reference to the accompanying drawings in which:—
-
FIG. 1 shows a gas turbine engine having a pressure-rise combustor and nozzle guide vane arrangement according to the present invention. -
FIG. 2 is an enlarged schematic of a pressure-rise combustor and nozzle guide vane arrangement according to the present invention. -
FIG. 3 is an enlarged schematic of a further pressure-rise combustor and nozzle guide vane arrangement according to the present invention. - A
gas turbine engine 10, as shown inFIG. 1 , comprises in axial flow series aninlet 12, alow pressure compressor 14, ahigh pressure compressor 16, acombustion section 18, ahigh pressure turbine 20, alow pressure turbine 22 and anexhaust 24. Thelow pressure turbine 22 is arranged to drive thelow pressure compressor 14 via ashaft 26 and thehigh pressure turbine 20 is arranged to drive thehigh pressure compressor 16 via ashaft 28. Thecombustion section 18 comprises at least one pressure-rise combustor 30 and the at least one pressure-rise combustor 30 is positioned upstream of a stage of turbinenozzle guide vanes 32. Thegas turbine engine 10 is generally conventional and it operates in a conventional manner. - The
vanes 33 of the stage of turbine nozzle guide vanes 32 form an ejector, shown more clearly inFIG. 2 , wherein each of thevanes 33 of the stage of turbinenozzle guide vanes 32 has anupstream portion 34 and adownstream portion 36. Theupstream portions 34 of thevanes 33 have leadingedges 38 and theupstream portions 34 of thevanes 33 are arranged to extend substantially in straight lines. Theupstream portions 34 of thevanes 33 are arranged substantially parallel to each other to define constantarea mixing passages 40 for a flow of gases there-through. Thedownstream portions 36 of thevanes 33 are arranged at an angle α to theupstream portions 34 of thevanes 33 to turn the flow of gases there-through. Thus thelow pressure compressor 14 andhigh pressure compressor 16 are upstream of the at least one pressure-rise combustor 30 and thehigh pressure turbine 20 andlow pressure turbine 22 are downstream of the stage of turbinenozzle guide vanes 32. - The at least one pressure-
rise combustor 30 may be a pulsejet combustor, a valve-less pulsejet combustor or a pulse detonation combustor. - Preferably there are a plurality of pressure-
rise combustors 30 positioned upstream of the stage of turbinenozzle guide vanes 32. The pressure-rise combustors 30 are arranged circumferentially around an axis X-X of rotation of thegas turbine engine 10. - The pressure-
rise combustors 30 haveinlets 42 at their upstream end to receive air from thehigh pressure compressor 16 andoutlets 44 at their downstream ends to discharge exhaust gases towards the turbine nozzle guide vanes 32 andhigh pressure turbine 20. Theoutlets 44 of the pressure-rise combustors 30 have a diameter D and are spaced from the leadingedges 38 of the turbinenozzle guide vanes 32 by a distance L. - Preferably the ratio of the distance L between the
outlets 44 of the at least one pressure-rise combustor 30 and the leadingedges 38 of theupstream portions 34 of thevanes 33 of the turbinenozzle guide vanes 32 to the diameter D of theoutlet 44 of the at least one pressure-rise combustor 30 is 1 to 1 or less than 1 to 1. The turbine nozzle guide vanes 32 have a contraction ratio in the range of 8 to 1 to 30 to 1. The contraction ratio of the turbinenozzle guide vanes 32 is the ratio of flow area AI into the turbinenozzle guide vanes 32 to the flow area AO out of the turbinenozzle guide vanes 32. - The distance L from the
outlet 44 of the pressure-rise combustor 30 to the leadingedge 38 of thevanes 32 is maintained small to retain maximum mechanical energy in the flow of exhaust gases from the pressure-rise combustors 30. - In another embodiment of the present invention it may be possible to reduce the number of pressure-
rise combustors 30, but to increase the size of the pressure-rise combustors 30 and this increases the efficiency, or pressure-rise/pressure-gain, generated by the pressure-rise combustor 30. This would require the exhaust gases from each pressure-rise combustor 30 to flow into a plurality of mixingpassages 40 between theupstream portions 34 ofadjacent vanes 32 and one ormore splitter vanes 46 would be required in themixing passages 40 between thedownstream portions 36 ofadjacent vanes 32. Thus, in this arrangement theupstream portions 34 of thevanes 33 again define constantarea mixing passages 40 and each constantarea mixing passage 40 supplies a flow of gases intopassages 48 between thedownstream portions 36 of thevanes 33 and thesplitter vanes 46. Thesplitter vanes 46 are located equi-distant between thedownstream portions 36 of thevanes 33 and thesplitter vanes 46 have substantially the same profile as thedownstream portions 36 of thevanes 33. The leadingedges 50 of thesplitter vanes 46 may extend in an upstream direction slightly into themixing passage 40 between theupstream portions 34 of thevanes 33. If there is more than onesplitter vane 46 betweenadjacent vanes 33, then some of thepassages 48 are defined between twosplitter vanes 46. In this case the distances between adjacent splitter vanes 46 and the distances betweensplitter vanes 46 and thedownstream portions 36 of thevanes 33 are equal. - Furthermore, the ratio of the number of splitter vanes to pressure-
rise combustors 30 may be increased by providing a plurality ofsplitter vanes 46 between thedownstream portions 36 ofadjacent vanes 32. The leadingedge 38 of theupstream portions 34 of thevanes 32 may have a part-circular cross-section or a square cross-section. - In the embodiments mentioned above the cross-sectional area of the
outlet 44 of the pressure-rise combustor 30 is less that the cross-sectional area of the inlet to themixing ducts 40 and preferably the ratio of the cross-sectional area of the inlet to the mixing passage to the cross-sectional area of the outlet of the pressure-rise combustor is 5.5 to 1. - Adjacent pressure-rise combustors may be arranged to fire in phase or out of phase.
- The present invention provides an axially compact arrangement for a pressure-rise, or pressure-gain, combustor compared with the previous arrangements with an ejector in flow series between the pressure-rise combustor and a stage of turbine nozzle guide vanes. The present invention enables the axial length of a gas turbine engine having a pressure-rise combustor to be reduced, reducing the axial length of shafts and hence the weight and cost of the gas turbine engine. In addition the present invention reduces the amount of material exposed to high temperature exhaust gases and thus reducing the cooling requirements. Also the exhaust gases from the pressure-rise combustor are mixed with the ambient gas around the pressure-rise combustor in an accelerating region within the passages between the vanes of the stage of turbine nozzle guide vanes and therefore entropy production, or irreversibilities, are minimised, thereby increasing the efficiency of the pressure-rise combustor. The present invention produces a steady flow of gases at the outlet of the turbine nozzle guide vanes from the unsteady flow of exhaust gases from the outlets of the pressure-rise combustors.
- The gas turbine engine may be a turbofan gas turbine engine, a turbopropeller gas turbine engine, a turboshaft gas turbine engine or a turbojet gas turbine engine.
Claims (11)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GBGB0822676.3A GB0822676D0 (en) | 2008-12-12 | 2008-12-12 | A gas turbine engine |
GB0822676.3 | 2008-12-12 |
Publications (2)
Publication Number | Publication Date |
---|---|
US20100150705A1 true US20100150705A1 (en) | 2010-06-17 |
US8413418B2 US8413418B2 (en) | 2013-04-09 |
Family
ID=40325996
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/591,260 Expired - Fee Related US8413418B2 (en) | 2008-12-12 | 2009-11-13 | Gas turbine engine |
Country Status (3)
Country | Link |
---|---|
US (1) | US8413418B2 (en) |
EP (1) | EP2196626A3 (en) |
GB (1) | GB0822676D0 (en) |
Cited By (9)
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US8571723B2 (en) | 2011-12-28 | 2013-10-29 | General Electric Company | Methods and systems for energy management within a transportation network |
US8655518B2 (en) | 2011-12-06 | 2014-02-18 | General Electric Company | Transportation network scheduling system and method |
WO2014055104A1 (en) * | 2012-10-01 | 2014-04-10 | United Technologies Corporation | Gas turbine engine with first turbine vane clocking |
US8805605B2 (en) | 2011-05-09 | 2014-08-12 | General Electric Company | Scheduling system and method for a transportation network |
US8818584B2 (en) | 2011-12-05 | 2014-08-26 | General Electric Company | System and method for modifying schedules of vehicles |
US9008933B2 (en) | 2011-05-09 | 2015-04-14 | General Electric Company | Off-board scheduling system and method for adjusting a movement plan of a transportation network |
US9156483B2 (en) | 2011-11-03 | 2015-10-13 | General Electric Company | System and method for changing when a vehicle enters a vehicle yard |
JP2015531456A (en) * | 2012-10-12 | 2015-11-02 | キング アブドラ ユニバーシティ オブ サイエンス アンド テクノロジー | Static detonation wave engine |
US9235991B2 (en) | 2011-12-06 | 2016-01-12 | General Electric Company | Transportation network scheduling system and method |
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WO2013139404A1 (en) | 2012-03-23 | 2013-09-26 | Institut Fuer Luftfahrtantriebe (Ila) Universitaet Stuttgart | Blade row for an unsteady axial flow gas turbine stage |
US9869190B2 (en) | 2014-05-30 | 2018-01-16 | General Electric Company | Variable-pitch rotor with remote counterweights |
US10072510B2 (en) | 2014-11-21 | 2018-09-11 | General Electric Company | Variable pitch fan for gas turbine engine and method of assembling the same |
US10100653B2 (en) | 2015-10-08 | 2018-10-16 | General Electric Company | Variable pitch fan blade retention system |
FR3095239B1 (en) | 2019-04-18 | 2021-05-07 | Safran | Transition piece for a gas turbine |
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2008
- 2008-12-12 GB GBGB0822676.3A patent/GB0822676D0/en not_active Ceased
-
2009
- 2009-11-13 EP EP09252615.1A patent/EP2196626A3/en not_active Withdrawn
- 2009-11-13 US US12/591,260 patent/US8413418B2/en not_active Expired - Fee Related
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Cited By (10)
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US8805605B2 (en) | 2011-05-09 | 2014-08-12 | General Electric Company | Scheduling system and method for a transportation network |
US9008933B2 (en) | 2011-05-09 | 2015-04-14 | General Electric Company | Off-board scheduling system and method for adjusting a movement plan of a transportation network |
US9156483B2 (en) | 2011-11-03 | 2015-10-13 | General Electric Company | System and method for changing when a vehicle enters a vehicle yard |
US8818584B2 (en) | 2011-12-05 | 2014-08-26 | General Electric Company | System and method for modifying schedules of vehicles |
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Also Published As
Publication number | Publication date |
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GB0822676D0 (en) | 2009-01-21 |
EP2196626A3 (en) | 2014-12-17 |
EP2196626A2 (en) | 2010-06-16 |
US8413418B2 (en) | 2013-04-09 |
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